This invention relates generally to turbomachinery, and, more specifically, to methods and systems for providing cooling for internal structures of gas turbine components.
In at least some known gas turbines, an internal structure of a component that is exposed to hot combustion gases is cooled using a cooling fluid that is channeled through passages defined within the component. In components such as stator vanes and rotor blades that extend substantially radially with respect to an axis of a gas turbine, at least some of the cooling passages likewise extend substantially radially. At least some further passages extend below and substantially parallel to at least a portion of an outer surface of the component. Cooling fluid is supplied to the passages from a source of cooling fluid coupled to the component.
Moreover, in at least some known gas turbines that include multiple rotor and stator stages, trailing-edge areas of airfoils of first-stage stator nozzle vanes, and first-stage rotor blades as well, experience temperatures and corresponding thermal loads that are amongst the highest that are encountered within a gas turbine. Accordingly, there is a tendency for a designer to increase a thickness of an airfoil, to provide a structural volume that is sufficiently large to facilitate defining cooling passages therein. However, there is a competing pressure on designers to reduce airfoil thickness, particularly in the trailing-edge areas, as trailing-edge thickness is a factor that exerts significant influence on aerodynamic efficiency of an airfoil.
Accordingly, it is desirable to improve airfoil aerodynamic efficiency by reducing trailing-edge thickness, while simultaneously facilitating enhanced cooling of trailing-edge structures.